Conventionally, a bypass turbojet 10, as shown in FIG. 1, is constituted by a gas turbine 12 of axis 14 driving a ducted fan 16 which is generally located upstream (UP) from the engine. The mass of air sucked in by the engine is split into a primary air stream (arrow A) that flows through the gas turbine or engine core, and a secondary air stream (arrow B) that comes from the fan 16 and that surrounds the engine core, the primary and secondary air streams being coaxial.
In well-known manner, the primary air stream (arrow A) is generally compressed by a first compressor 18 referred to as a low pressure (LP) compressor or booster, having an LP shaft that is connected to the shaft of the fan 14 and that is driven in rotation by the shaft of the downstream low pressure turbine (not shown), the air then being compressed in a second compressor 20 further downstream (DN), referred to as a high pressure (HP) compressor, having an HP shaft that is driven in rotation by the shaft of a high pressure turbine arranged at the outlet from a combustion chamber and located upstream from the low pressure turbine (the combustion chamber and the turbines not being shown).
In such a two-spool turbojet, the term “intermediate casing” 22 is commonly used to designate a casing having its hub arranged between a casing 24 of the low pressure compressor 18 and a casing 26 of the high pressure compressor 20.
The intermediate casing 22 has an inner annular wall 28 defining the outside of the annular primary air flow passage 18, an intermediate annular wall 30 defining the inside of the annular secondary air flow passage 33, and an outer wall 35 defining the outside of the annular secondary air flow passage 33.
Furthermore, such a turbojet is generally provided with devices known as variable bleed valves (VBVs) 32 that serve to divert a portion of the primary air stream at the outlet from the LP compressor 18 into the annular channel 33 of the secondary air stream. By lowering the pressure downstream from the LP compressor 18, this bleeding has the effect of lowering its operating point and thus of reducing the risks of the compressor 18, 20 surging which would lead to a sudden reversal of the flow direction of the hot gas stream from the combustion chamber, and which could damage the compressor 18, 20. Furthermore, in the event of accidental penetration of water, in particular in the form of rain or hail, or indeed in the event of accidental penetration of various kinds of debris, which can harm the operation of a turbojet, these valves make it possible to recover such water or debris and eject it from the primary passage feeding air to the combustion chamber.
Thus, the bleed valves 32 are formed in the inner annular shroud 28 of the hub of the intermediate casing 32 and they communicate with a space lying between the inner annular shroud 28 and the intermediate shroud 30 of the intermediate casing 22.
In order to bleed off air, the hub of the intermediate casing 22 has a downstream transverse plate 34 arranged upstream from the high pressure compressor 20 of the turbojet and connecting together the downstream ends of the inner and intermediate annular shrouds 28 and 30. The downstream plate 34 has a plurality of first openings 36 distributed around the axis 14 of the turbojet 10, each communicating upstream with the inside of the hub and downstream with a duct 38 having its downstream end leading to a shroud 40 that is perforated by second openings downstream from an outer annular shroud 42 formed to extend the intermediate annular wall 30 of the intermediate casing 32 downstream.
As shown in FIG. 1, the hub of the intermediate casing 22 carries stator vanes 44 that extend between the intermediate wall 30 and the outer wall 35 of the intermediate casing 22. The stator vanes 44, also known as outlet guide vanes (OGVs), are for straightening out the secondary air stream coming from the upstream fan 16.
In the context of developing a turbine engine, it is necessary to measure and verify its performance. It is desired in particular to measure the flow parameters of the stream flowing in the secondary passage, such as its speed, its pressure, and its temperature. For this purpose, it has been found that it is preferable to arrange measurement sensors at certain precise locations in the secondary passage. One of these locations is situated downstream from the stator vanes 44 of the intermediate casing 22, in a plane lying at a particular angle of inclination relative to the axis of the turbine engine and passing via the perforated shroud 40. This location makes it possible to take good measurements of the performance of the assembly comprising the fan 16 and the guide vanes 44. In order to take exhaustive measurements of the stream in this location, it is desirable to arrange a plurality of sensors at different heights in the secondary passage, while remaining in this plane. Even though these sensors are incorporated in intrusive manner, they must not influence the normal operation of the turbine engine, and they must be capable of withstanding the environment in which they are to be found during testing, where such environments generally cover all of the possible operating ranges of the turbine engine. In particular, during such tests, it is possible to observe variations in temperature, in pressure, and in relative positioning of parts because of various assembly clearances and because of differential expansions.